The present disclosure relates generally to gas turbine engines and, more specifically, to systems and methods of forming protective caps around cooling holes in a gas turbine component.
In a gas turbine engine, air pressurized in a compressor is mixed with fuel in a combustor to generate hot combustion gases. Energy is initially extracted from the gases in a high pressure turbine (HPT) that powers the compressor, and subsequently in a low pressure turbine (LPT) that powers a fan in a turbofan aircraft engine application, or powers an external shaft for marine and/or industrial applications. Generally, engine efficiency increases as the temperature of combustion gases is increased. However, the increased gas temperature increases the operating temperature of various components along the gas flowpath, which in turn increases the need for cooling such components to facilitate extending their useful life.
For example, at least some known gas turbine components, such as blades, nozzles, and liners, require cooling during operation of the gas turbine engine. In at least some gas turbine engines, flowpath components exposed to hot combustion gases are cooled using compressor bleed air. For example, at least some known components channel the compressor bleed air through film cooling holes defined within the gas turbine components. However, the gas turbine components generally have a limited service life and must be periodically serviced to ensure the gas turbine components continue to function properly. Servicing the gas turbine components typically includes removal of an existing thermal barrier coating and subsequent reapplication of a thermal barrier coating to the components. The film cooling holes may become blocked when reapplying the thermal barrier coating, and cleaning and clearing the film cooling holes of the coating is a time-consuming and laborious task.